The present invention relates to gas turbine engine blades. In particular, the present invention relates to the internal cooling configuration of turbine blades.
A gas turbine engine commonly includes a fan, a compressor, a combustor, a turbine, and an exhaust nozzle. During engine operation, working medium gases, for example air, are drawn into and compressed in the compressor. The compressed air is channeled to the combustor where fuel is added to the air and the air/fuel mixture is ignited. The products of combustion are discharged to the turbine section, which extracts a portion of the energy from these products to power the fan and compressor. The fan and compressor together with the energy in the combustion products not used by the turbine to drive the fan and compressor produce useful thrust to power, for example, an aircraft in flight.
The compressor and turbine commonly include alternating stages of rotor blades and stator vanes. Compressor and turbine blades and vanes often include complex, contoured blade geometries designed to optimally interact with the working medium gas passing through the engine. Additionally, the operating temperatures of some engine stages, such as in the high pressure turbine stages, may exceed the material limits of the blades and therefore necessitate cooling the blades. Cooled blades may include cooling channels, sometimes referred to as passages, in various configurations through which a coolant, such as compressor bleed air, is directed to convectively cool the blade. Blade cooling channels may be oriented spanwise from the root to the tip of the blade or axially between leading and trailing edges. The channels may be fed by one or more supply channels located toward the root, where the coolant flows radially outward from the root to tip, in what is sometimes referred to as an “up-pass.” Alternatively, the channels may be fed by one or more supply channels located toward the tip of the blade, in a so-called “down-pass.” In addition to individual up and down passes, some blades include cooling channels in a serpentine configuration consisting of several adjacent up and down-passes proceeding axially forward or afterward through the blade. The blades may also include other cooling features, such as film cooling holes for exhausting the coolant from the cooling channels over the exterior surface of the blade, as well as impingement cooling walls, trip strips, and turbulators.
Prior turbine blade designs have continually sought to decrease blade temperatures through cooling. A particular challenge in prior cooled blades lies in cooling the concave pressure side of turbine blades where temperatures are generally higher than, for example, the convex suction side of the blade. Prior cooled blades and vanes have employed several techniques to cool the pressure side including internal cooling channels supplied by bleed air from the compressor and adapted to eject the cooling fluid through film cooling holes onto the exterior of the pressure side wall. However, due to the relatively high temperatures encountered during operation, a need still exists to improve the preferential cooling of turbine blade and vane airfoil pressure side walls.